For gas turbine engines, it is generally known that the operational clearances between the tips of rotating blades and engine static structure impact the thermodynamic efficiency and fuel burn of the engine. Hence, gas turbine engine manufacturers continually seek ways to reduce these operational clearances. The value of even several thousandths of an inch improvement can be quite significant, especially in the high pressure turbine and high pressure compressor. As a result, many gas turbine engine manufacturers trade markedly higher manufacturing costs in exchange for small improvements in blade tip clearance. These costs can be embedded in complex design features, in high precision manufacturing tolerances, and exotic build processes as a means to achieve reduced blade tip clearance. Despite such efforts, typically two to five thousandths of an inch in tip clearance is needed to accommodate geometric uncertainty in the location of the rotor centerline with respect to key locations on the static structure.
In addition to the operational clearances described above, gas turbine engine rotor dynamics receive great attention during engine design. This includes the placement of shaft critical speed in the frequency domain, and the rotor response to imbalance and transient excursions through critical speeds. Critical speed placement is controlled primarily via stiffness in the rotor/bearing support, while rotor response to imbalance and transient critical speed operation is controlled via damping. Typically, damping and stiffness control are provided via hydraulic devices, such as “squeeze film dampers” (SFDs), at rotor bearing locations. As is generally known, SFDs achieve both stiffness and damping via the whirl motion of the shaft within a controlled oil film annulus. However, both the stiffness and the damping coefficient achieved are highly non-linear with respect to orbital (whirl) displacement of the shaft. Moreover, the stiffness and damping coefficients are inexorably linked, which means one cannot be modified without a large effect on the other. This results in an inability to precisely locate and control response to critical speeds, since stiffness and damping are varied along with whirl displacement. This variability and imprecision causes manufacturers to design gas turbine engines with substantial frequency margin above running speeds for shaft bending mode critical speeds, and with having to accept some uncertainty in the placement and response of rigid rotor modes, which are commonly traversed in transient speeds during start and shutdown.
The net effect of the tip clearance and shaft dynamics issues described above can result in reduced efficiency and increased product cost, with additional costs embedded in a reduced yield in the assembly/test process due to the incidences of engines failing to meet specifications for temperature or vibration.
Hence, there is a need for a rotor tip clearance and shaft dynamics system and methods for gas turbine engines that provides increased efficiency and reduced operational and manufacturing costs. The present invention addresses at least this need.